1. Field of the Invention
The present invention relates to a gauge for measuring heat fluxes, especially in high temperature, highly corrosive environments. In particular, the present invention relates to a gauge for measuring heat fluxes in rocket motor nozzles, including sub-scale and full-scale rocket motor nozzles, heat shields of re-entry vehicles, rocket motor combustion chambers, and metal processing. This invention also relates to a method of measuring heat fluxes with the gauge in various high temperature, corrosive environments.
2. Description of the Related Art
Solid rocket motors typically include an outer case or shell housing a solid propellant grain which, in the case of a hybrid motor, is a solid fuel or oxidizer grain. The rocket motor case is conventionally manufactured from a rigid, yet durable, material such as steel or filament-wound composite. The propellant is housed within the case. There are several basic propellant grain configurations known in the art (and compatible with the use of this invention) for loading propellant within the case. The two most commonly used configurations are the center-perforated grain configuration and the end burning grain configuration. In the center-perforated grain configuration, the flame front advances radially from the center perforation towards the outer case. On the other hand, in the end-burning grain configuration the flame front advances axially from the nozzle end of the motor towards the forward dome.
During firing, oxidizing agents of the propellant serve to drive combustion reactions in a combustion chamber to form large quantities of combustion products, which are expelled from the rocket motor through a nozzle in fluid communication with the combustion chamber. Nozzles are designed to accelerate the combustion product gases from the propellant grain to the maximum velocity at the exit. To achieve this end, nozzles usually have forward walls converging to a restricted throat region and aft walls diverging from the throat region to a larger exit area, thus defining a converging/diverging contoured pathway. The nature of compressible gases is such that a converging/diverging nozzle increases the exit gas velocity and thereby thrust. The proportions of the mass flow pathway, particularly the ratio of area at the exit plane to area at the throat, establish how efficiently the nozzle converts pressure in the mass flow stream to thrust produced by the motor. It is within the purview of those skilled in the art to design a nozzle throat to optimize the ratio of exit area to throat area.
During operation, combustion of solid rocket propellant generates extreme conditions within the case and along the contoured nozzle pathway of the rocket motor. For example, temperatures inside the rocket motor case can exceed 2760xc2x0 C. (5,000xc2x0 F.), and interior pressures can exceed 1,500 psi. These factors combine to create a high degree of turbulence for particles entrained in the combustion gases.
A heat insulating layer (insulation) protects the rocket motor case from the hot gas and highly erosive particle streams generated by the combusting propellant. Typically, the propellant grain is bonded to the insulation and/or non-insulated portions of the case by use of a lining layer (liner). In addition to its adhesive function of bonding the propellant to the insulation and any non-insulated portions of the case, the liner also supplements the insulator by functioning to inhibit the burning surface of the propellant grain when the propellant/liner interface is exposed to an approaching flame front. Additionally, the liner isolates propellant within the case from the environment and prevents leakage of combustion gases or liquid into or through the case.
Likewise, the contoured nozzle pathway, including the restricted nozzle region, must also be insulated to withstand the elevated temperatures and pressures of the combustion products, as well as the erosive effects of turbulent particles entrained within the combustion gas. Carbon-based and silica-based ablative materials are highly advantageous for use as nozzle insulation due to their excellent ablative properties, low cost, and relatively light weight.
It is widely acknowledged in the industry, however, that carbon-and silica-based insulation and ablative materials, such as those present at case insulation-to-propellant interfaces and those defining nozzle contour pathways, are highly susceptible to recession at high operating temperatures. Convective and radiative heating of the ablative materials by the combustion products increases the vulnerability of the ablative materials to recession. As incident heat is conducted to the nozzle throat ablative material, the ablative material tends to decompose into pyrolysis gases and residual carbon, which are carried away with the propellant combustion gases. If not accounted for, the recession of the nozzle throat inner surface during motor operation may become a source of several problems in rocket operation, including decreased efficiency and loss of predictability.
Thus, the design of rocket insulation, including the selection of an appropriate ablative materials and insulation thickness, is dependent upon the convective and radiative heat fluxes incident on ablative surfaces at various locals in rocket motors and nozzles. Because no reliable measuring technique has heretofore been known for determining actual heat fluxes, heat fluxes are usually estimated by modeling based on data obtained from such techniques as computational fluid dynamics.
The criticality of predication accuracy is demonstrated by the severity and magnitude of the risk of failure due to erosion. Most insulation is, of necessity, xe2x80x9cman-ratedxe2x80x9d in the sense that a catastrophic failure can result in the loss of human lifexe2x80x94whether the rocket motor is used as a booster for launch of a manned rocket or is carried tactically underneath the wing of an attack aircraft. The monetary cost of failure in satellite launches is well-publicized and can run into the hundreds of millions of dollars. Additionally, as mentioned above, unforeseen amounts of recession in the nozzle throat insulation can significantly affect the flow expansion contour in the nozzle, affecting motor performance.
For these reasons, there is a strong desire in the art to validate heat flux predictions through actual measurement of the convective and radiative heat fluxes incident on the ablative surfaces of rocket motors and nozzles. As described in Wool et al., xe2x80x9cMeasurement of Convective and Radiative Heat Fluxes at the Surface of an Ablative Materialxe2x80x9d (1970), thermocouples have been proposed for validating heat flux predictions. Wool et al. state that thermocouples used for this purpose should be capable of accurately measuring the high temperatures encountered in rocket motor materials and, because ablative material recedes and decomposes, should be capable of obtaining sufficient data for a continuous evaluation of the heat flux. Additionally, the gauge must be capable of withstanding the high operating temperatures to which it is exposed for a sufficient time to obtain heat flux data. Although the gauges disclosed in Wool et al. reportedly attain these objects, the structure of the Wool et al. thermocouple devices requires that data reduction be done by a complex analytical technique in which assumptions are made concerning phenomena such as surface thermochemical reactions, transient heat conduction, and in-depth pyrolysis gas generation.
It is, therefore, an object of this invention to provide a gauge capable of measuring either heat fluxes (convective and radiative fluxes) or radiative fluxes at a rocket motor ablative surface with sufficiently high accuracy and for a sufficiently long period of time to permit validation of known prediction models, without the need for complex analytical data reduction.
In accordance with the principles of this invention, the above and other objects are attained by a novel gauge for measuring heat flux, especially heat flux encountered in a high temperature environment, such as a rocket motor. The gauge comprises at least one thermocouple and an anisotropic pyrolytic graphite body that covers at least part of, and optionally encases, the thermocouple. The anisotropic pyrolytic graphite provides a body on which heat flux is incident, so that changes in temperature to the body can be measured by the thermocouple. These measurements can then be reduced to quantify the heat flux incident to the body. In the event that the thermocouple is encased in an anisotropic pyrolytic graphite body, a bore may be formed in the body for permitting the thermocouple to be inserted into the body and for providing a pathway for lead wires of the thermocouple. The bore can be filled with, for example, boron nitride.
Pyrolytic graphite has a series of basal planes, in which carbon atoms are arranged in repeating hexagonal patterns. The direction in which the planes extend is commonly referred to as the xe2x80x9caxe2x80x9d direction or the xe2x80x9ca-bxe2x80x9d plane. Individual basal planes are stacked one upon another along a direction perpendicular to the basal planes which is commonly known as the xe2x80x9ccxe2x80x9d direction. Unlike single-crystal graphite in which the planes are ordered, the planes are arranged somewhat randomly for pyrolytic graphite.
Pyrolytic graphite is characterized as anisotropic because the properties of pyrolytic graphite along the xe2x80x9caxe2x80x9d direction differ greatly from the properties of pyrolytic graphite along the xe2x80x9ccxe2x80x9d direction. For example, pyrolytic graphite exhibits high tensile strength, low thermal expansion, and, most importantly for the present invention, high thermal conductivity in the xe2x80x9caxe2x80x9d direction along the basal planes. On the other hand, the thermal conductivity of pyrolytic graphite in the xe2x80x9ccxe2x80x9d direction has been reported to be about 0.59 Btu/ft/hr/xc2x0 F. (at 2000xc2x0 F.), which is less than 1% of the reported thermal conductivity of 114 Btu/ft/hr/xc2x0 F. (at 2000xc2x0 F.) for pyrolytic graphite in the xe2x80x9caxe2x80x9d direction.
By arranging the gauge so that the gauge surface on which convective and radiative fluxes are incident is perpendicular to the basal planes of the pyrolytic graphite, the conductivity of the pyrolytic graphite permits energy, transferred into the pyrolytic graphite body in the form of heat flux on the incident (or facing) surface, to be quickly distributed through the entire pyrolytic graphite body, resulting in small substantially instantaneous temperature gradients. As a result, measurements of the rate of temperature change are simplified, no matter at what depth (or height) in the gauge that the thermocouple is disposed, since the rate of temperature change will be substantially uniform throughout the entire pyrolytic graphite body.
The gauge of this invention is especially useful for permitting measurement of heat fluxes along nozzle surfaces defining contoured pathways, such as diverging/converging pathways. The gauge can also be used in other regions of a rocket motor, such as the aft dome, combustion chamber, or interface between the case insulation and propellant grain. It is also within the scope of this invention to use the gauge for measuring rates of temperature changes caused by heat fluxes at heat shields of reentry vehicles or other regions of vehicles exposed to high heat flux. The gauge also finds applicability in other technologies, including, for example, metal processing.
Other objects, aspects, and advantages of this invention will become more apparent to those skilled in the art upon reading the specification and appended claims which, when taken in conjunction with the accompany drawing, explain the principles of this invention.